1. Field of the Invention
The present invention is related to a gas turbine, and to a gas bleeding method for a gas turbine, which perform sealing between moving blades and stationary blades by supplying bleed gas from, for example, a compressor, while cooling the moving blades.
2. Description of the Related Art
In a gas turbine plant, compressed air from a compressor is fed to a combustor, wherein it is combusted along with fuel to generate high temperature gas, which is conducted to a gas turbine so as to drive said gas turbine. And there is a per se known structure in which, at this time, a portion of this compressed air is conducted as bleed gas to a cooling device, and after being cooled this bleed gas is next fed to stationary blades and moving blades on the gas turbine side, so that this bleed gas is utilized for cooling of these moving blades and secondary blades, and for sealing between these moving blades and secondary blades. An example of a structure in such a prior art gas turbine for supplying bleed gas to the stationary blades and the moving blades of a first stage unit will now be described in the following with reference to FIG. 3. This figure is a partial axial cross sectional view showing a bleed gas flow conduit to the first stage unit of the gas turbine, and it should be understood that a compressor which is not shown in the drawing and lies beyond the extreme left margin of the drawing paper disposed coaxially with the gas turbine.
In this figure, the reference numeral 1 indicates a set of first stage moving blades, while the reference numeral 2 indicates a set of first stage stationary blades. A plurality of first stage moving blades 1 are disposed in circular arrangement around the periphery of a rotor disk 3 which is mounted coaxially with the compressor, and this first stage rotor disk 3 rotates by receiving the impulse of combustion gas from said compressor. Furthermore, a plurality of first stage stationary blades 2 are disposed in a circular arrangement so as to be coaxial with the first stage rotor disk 3, on near side of the turbine casing. Thus a first stage unit 4 is constituted, comprising these first stage moving blades 1, this first stage rotor disk 3, and this first stage stationary blades 2.
Furthermore, the reference numeral 5 in the figure indicates a bleed gas chamber which takes in a flow f1 of bleed gas from the previously described cooler after said bleed gas flow has been cooled, and almost all of this bleed gas flow f1 which has been taken into the bleed gas chamber 5 is conducted to the first stage moving blades 1 via a cooling flow conduit 3a which is formed in the first stage rotor disk 3, and thus functions to cool these first stage moving blades 1 from their insides. That is, the cooling flow conduit 3a is a flow conduit which is formed in roughly an xe2x80x9cLxe2x80x9d shape between the upstream side surface of the first stage rotor disk main body 3b (the surface thereof which confronts the first stage stationary blades 2) and a flow conduit partition wall 3c which is fixed by bolts to said upstream side surface; and, after a cooling air flow f2 has been taken in along the direction of the rotational axis of the first stage rotor disk 3 from the bleed gas flow f1 being expelled from the bleed gas chamber 5, next this cooling air flow f2 is expelled along the radial direction with respect to said rotational axis as a center.
This flow conduit partition wall 3c is a tubular member which partitions the flow f1 of bleed gas from the bleed gas chamber 5 into two flows, the aforesaid cooling air flow f2 and a sealing air flow f3; and a labyrinth seal 6 is formed upon its outer circumferential surface, between the flow conduit partition wall 3c and a division wall 2a1 which is held by the inner circumferential side of an inner shroud 2a of the first stage secondary blades 2.
A portion of the bleed gas flow f1 is separated to constitute said sealing air flow f3, which is then supplied between the first stage moving blades 1 and the first stage secondary blades 2; and this labyrinth seal 6 functions to seal these gaps C.
However, such a prior art type gas turbine suffers from the problems explained below. That is, the bleed gas flow f1 which is supplied from the bleed gas chamber 5 has hardly any rotational speed component around the circumferential direction of said rotational axis taken as a center, and, since it enters into the disk holes 3a1 which are formed in the cooling flow conduit 3a (a plurality of perforations which are formed so as to radiate from said rotational axis) in this same state, there is the problem of occurrence of drive power loss.
That is, although the each cooling flow conduit 3a rotates at high speed together with the first stage rotor disk 3 which is the main rotating body, since the cooling air flow f2 which has hardly any high rotational velocity component in the circumferential direction with respect to the first stage rotor disks 3 in this high speed rotating state flows in and passes through the first stage for disk 3, accordingly this flow of cooling air f2 undesirably exerts a braking force to restrain the rotational operation of the first stage rotor disk 3; and, moreover, the drive power required for rotating the rotating body which includes the first stage rotor disk 3 is undesirably increased. It is desirable to eliminate the rotational power loss by all means possible, since this type of drive loss entails an undesirable reduction in the electric generating capacity of a generator (not shown in the figures) which is connected to the gas turbine.
The present invention has been made in consideration of the above described problems, and its objective is to provide a gas turbine and a gas bleeding method therefor, which are capable of preventing loss of drive power due to gas bleeding to the rotor disk.
The present invention utilizes the following means for solving the problems detailed above.
Namely, the gas turbine described in a first aspect of the present invention comprises a plurality of stationary blades arranged in a circular manner on near side of a turbine casing, a plurality of moving blades arranged in circular manner on near side of a rotor disk adjoining the stationary blades, a swirling flow creation section which supplies to the rotor disk bleed gas which has been input, after imparting the bleed gas with a swirling flow which rotates in the same rotational direction as that of the rotor disk, and a seal gas supply flow conduit which supplies a portion of the bleed gas to a gap between the stationary blades and the moving blades, bypassing the swirling flow creation section.
According to the gas turbine specified in the first aspect of the present invention as described above, the flow of bleed gas is supplied towards the rotor disk after having been imparted with a swirling flow by passing through the swirling flow creation section, and therefore it becomes possible to greatly reduce the relative rotational speed difference between the two of them (the rotor disk and the bleed gas flow) in the rotational direction of the rotor disk. Moreover, the bleed gas flow for sealing between the stationary blades and the moving blades is arranged to flow within the seal gas supply flow conduit, thus not interfering with the above described swirling flow in the swirling flow creation section.
Furthermore, a gas turbine described in a second aspect of the present invention, the swirling flow creation section comprises a plurality of TOBI nozzles (Tangential OnBoard Injection Nozzle) which reduce the flow conduit cross sectional area while swirling from the outside in the radial direction towards the inside, around the rotational axis of the rotor disk as a center; and the seal gas supply flow conduit is formed so as to pass between the TOBI nozzles.
According to the gas turbine specified in the second aspect of the present invention as described above, it is possible to impart a swirling action to the flow of gas towards the rotor disk in a reliable manner. Furthermore, it becomes possible to supply the bleed gas for sealing to the gap between the stationary blades and the moving blades without hampering this swirling flow.
A gas bleeding method described in a third aspect of the present invention, in a bleeding method for gas turbine which comprises a plurality of stationary blades arranged in a circular manner on near side of a turbine casing, a plurality of moving blades arranged in a circular manner on near side of a rotor disk adjoining the stationary blades; and in this method, bleed gas is supplied to the rotor disk after being imparted with a swirling flow which rotates in the same rotational direction as that of the rotor disk; and a portion of the bleed gas is supplied between the stationary blades and the moving blades bypassing the swirling flow.
According to the gas bleeding method specified in the third aspect of the present invention as described above, since the flow of bleed gas is supplied towards the rotor disk after having been imparted with a swirling flow, it becomes possible to greatly reduce the relative rotational speed difference between the two of them in the rotational direction of the rotor disk. Moreover, the bleed gas flow for sealing between the stationary blades and the moving blades does not interfere with the above described swirling flow.